FIG. 1 shows a high-bypass gas turbine engine 10. The engine 10 comprises, in axial flow series, an air intake duct 11, an intake fan 12, a bypass duct 13, an intermediate pressure compressor 14, a high pressure compressor 16, a combustor 18, a high pressure turbine 20, an intermediate pressure turbine 22, a low pressure turbine 24 and an exhaust nozzle 25. The fan 12, compressors 14, 16 and turbines 20, 22, 24 all rotate about the major axis of the gas turbine engine 10 and so define the axial direction of gas turbine engine.
Air is drawn through the air intake duct 11 by the intake fan 12 where it is accelerated. The accelerated air then passes through a plurality of Outlet Guide Vanes (OGVs) 31. A significant portion of the airflow is discharged through the bypass duct 13 generating a corresponding portion of the engine 10 thrust. The remainder is drawn through the intermediate pressure compressor 14 into what is termed the core of the engine 10 where the air is compressed. A further stage of compression takes place in the high pressure compressor 16 before the air is mixed with fuel and burned in the combustor 18. The resulting hot working fluid is discharged through the high pressure turbine 20, the intermediate pressure turbine 22 and the low pressure turbine 24 in series where work is extracted from the working fluid. The work extracted drives the intake fan 12, the intermediate pressure compressor 14 and the high pressure compressor 16 via shafts 27, 28, 30. The working fluid, which has reduced in pressure and temperature, is then expelled through the exhaust nozzle 25 and generates the remaining portion of the engine 10 thrust. Respective low, intermediate and high pressure shafts 27, 28, 30 couple the fan 12 to the low pressure turbine 24, the intermediate pressure compressor 14 to the intermediate pressure turbine 22 and high pressure compressor 16 to the high pressure turbine 20 respectively.
The compressors 14, 16 comprise alternating rotors and stators, which compress the air in use. The intermediate compressor 14 comprises a first stator 32 provided at an inlet to the core, rear of the intake fan 12. This first stator is termed the “Engine Section Stator” (ESS) or “core inlet stator”. The ESS is prone to ice buildup in use due to the low temperatures sometimes experienced during engine operation, particularly at the leading edge. Such ice can reduce the aerodynamic performance of the ESS, and/or cause damage to downstream components such as rotors of the compressors 14, 16 where large pieces of ice breakoff. Consequently, aircraft gas turbine engines generally include anti-icing systems for the ESS 32 to prevent ice build-up, and to melt any ice that accumulates.
In a first known system, electrical heaters located within the ESS 32 are used to heat the ESS 32. The electrical power for this system must be obtained from generators (not shown) driven by the gas turbine itself. This results in additional electrical load on the engine generators, and hence increased fuel burn, as the engine must be operated at higher powers to provide sufficient torque to power the generators.
In a second known system, engine bleed air is used to provide ESS de-icing. FIG. 2 shows such a system. Bleed air is taken from the intermediate compressor 14 into a duct 34, and fed through the ESS 32, before being vented overboard. However, use of this compressed air is not thermodynamically efficient, since the compressed air is not utilised in the thermodynamic cycle of the engine 10. Furthermore, large pipes are required to deliver this air to the ESS. These pipes carry hot, high pressure air, and so must be relatively heavy to accommodate this. The hot air also presents a fire risk should a leak occur. Consequently, while this method may provide adequate de-icing, it also increases the weight, complexity and cost of the engine, while reducing fuel efficiency.
The present disclosure describes an ESS anti-icing arrangement which seeks to overcome some or all of the above problems.